With reference to FIG. 1, a ducted fan gas turbine engine generally indicated at 10 has a principal and rotational axis X-X. The engine comprises, in axial flow series, an air intake 11, a propulsive fan 12, an intermediate pressure compressor 13, a high-pressure compressor 14, combustion equipment 15, a high-pressure turbine 16, and intermediate pressure turbine 17, a low-pressure turbine 18 and a core engine exhaust nozzle 19. A nacelle 21 generally surrounds the engine 10 and defines the intake 11, a bypass duct 22 and a bypass exhaust nozzle 23.
The gas turbine engine 10 works in a conventional manner so that air entering the intake 11 is accelerated by the fan 12 to produce two air flows: a first air flow A into the intermediate pressure compressor 13 and a second air flow B which passes through the bypass duct 22 to provide propulsive thrust. The intermediate pressure compressor 13 compresses the air flow A directed into it before delivering that air to the high pressure compressor 14 where further compression takes place.
The compressed air exhausted from the high-pressure compressor 14 is directed into the combustion equipment 15 where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive the high, intermediate and low-pressure turbines 16, 17, 18 before being exhausted through the nozzle 19 to provide additional propulsive thrust. The high, intermediate and low-pressure turbines respectively drive the high and intermediate pressure compressors 14, 13 and the fan 12 by suitable interconnecting shafts.
Other gas turbine engines are known in the art. Such engines may have an alternative number of interconnecting shafts (e.g. two) and/or an alternative number of compressors and/or turbines. Further the engine may comprise a gearbox provided in the drive train from a turbine to a compressor and/or fan.
The turbines and compressors each include a series of stages arranged in axial flow series. In the case of a turbine, each stage consists of an annular arrangement or row of nozzle guide vanes, followed by a row of rotating turbine blades. FIG. 2 shows an isometric view of a circumferential portion of a typical single stage cooled turbine. The nozzle guide vanes are static components mounted to the engine casing and comprise an aerofoil portion 31, and radially inner and outer platforms 33. The nozzle guide vanes are shaped to swirl the gas flow in the direction of the turbine blade rotation to provide an optimum angle of incidence on the turbine blades and increase the tangential momentum of the gas flow.
The turbine blade rotor includes a plurality of blades peripherally mounted to a rotor disc which is rotatable about the principal axis of the gas turbine engine. Each blade includes a radially inner platform 34 and an aerofoil portion 32. In the arrangement shown, the blade is a shrouded blade meaning an outer platform or shroud is mounted to the radial tip of the blades, the shrouds of adjacent blades abutting one another to provide a full annulus on the radially outer of the gas path. A static seal segment 35 is located radially outside of the shroud with the two components acting in concert to provide an air seal and a preferential air path between the blades, rather than over them.
The turbine blades translate the circumferential flow leaving the nozzle guide vanes into rotation of the disc. The adjacent aerofoil portions of the blades define a gas path passage which provides steady acceleration of the flow up to the smallest flow area known as the throat. As will be appreciated, the turbine vanes and blades, particularly the earlier stages, are required to operate in an extremely hot environment with the rotational speeds on the blades creating significant centrifugal loading. Ensuring that the vanes and blades have the necessary aerodynamic performance, efficiency, cost, life and weight makes the turbine blades and vanes one of the most technically challenging areas of the gas turbine engine.
The transverse cross section of the blade and vanes are governed by the aerodynamic properties, the permitted stress, material and cooling passages located within the blade. FIG. 3 shows a generic transverse schematic section of an aerofoil portion 310 which may be that of a turbine blade or nozzle guide vane. The aerofoil 310 includes a leading edge 312 and a trailing edge 314 with pressure 316 and suction 318 surfaces extending therebetween. The axial dimension of the aerofoil is commonly referred to as the chord, whilst the radial length of the aerofoil, the span.
The pressure 316 and suction 318 surfaces are provided by respective pressure and suction walls. The interior of the aerofoil includes cooling passages 320 which are defined by the pressure and suction walls and webs which extend therebetween. The cooling air passages deliver air to the interior and exterior surfaces of the aerofoil. The exterior cooling is achieved via various cooling hole arrangements such as the film cooling holes shown in FIG. 2. As will be appreciated, the specific internal architecture and external shape of a vane or blade will be specific to a given engine and may vary considerably from those shown in FIGS. 2 and 3.
The nozzle guide vanes and turbine blades of current state of the art turbines are generally made by investment casting which allow for the integral formation of the internal cooling passages 320. Once cast, the blades undergo a number of processes to, for example, provide cooling holes, thermal barrier coatings and removal of extraneous materials and features which result from the casting process.
Despite careful control measures, the number and complex nature of the manufacturing steps can lead to considerable variation in the final components which can affect blade performance and lifting. Consequently, the blades are assessed at various stages of production and non-conforming parts are recycled or scrapped.
One criteria for assessing cast components is aerofoil shape and wall thickness which are typically measured and compared to a reference shape. Due to the highly specific and complex geometry of the aerofoils, the number of variations in the geometric shape and wall thickness of manufactured parts can vary tremendously and determining what is and is not acceptable is often difficult to assess on a case by case basis.
The present invention seeks to provide a method of producing gas turbine blades which have an improved performance consistency.
Although the introduction and following description is focussed predominantly of turbine aerofoils, namely those of vanes and blades, it is to be noted that the invention is applicable to any aerofoil or impeller or end wall thereof. Hence, the invention may be applied to compressors, propellers, turbines, fans etc. Further, the invention may find use in any component having a gas washed surface, such as a wing of an aircraft, an engine nacelle, or the shape of a marine impeller.